Methods and assemblies for attaching airfoils within a flow path

ABSTRACT

Flow path assemblies for gas turbine engines are provided. For example, a flow path assembly comprises an inner wall; a unitary outer wall; and a plurality of nozzle airfoils having an inner end radially opposite an outer end. The unitary outer wall defines a plurality of outer pockets each configured for receipt of the outer end of one of the nozzle airfoils, and the inner wall includes defines a plurality of inner pockets each configured for receipt of the inner end of one of the plurality of nozzle airfoils. A portion of each inner pocket is defined by a forward inner wall segment and an aft inner wall segment. In another embodiment, a flow path assembly comprises an inner wall defining a plurality of bayonet slots that each receive a bayonet included with each of a plurality of nozzle airfoils that are integral with a unitary outer wall.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of and claims priority to U.S.application Ser. No. 16/377,398 (issued as U.S. Pat. No. 11,286,799),filed Apr. 8, 2019, which is a continuation of and claims priority toU.S. application Ser. No. 15/440,294 (issued as U.S. Pat. No.10,253,641), filed Feb. 23, 2017, the contents of both of which areincorporated herein by reference.

FIELD

The present subject matter relates generally to gas turbine engines.More particularly, the present subject matter relates to outer and innerflow path boundary configurations for receipt of stator airfoils, aswell as methods for assembling stator airfoils to a gas turbine flowpath assembly.

BACKGROUND

A gas turbine engine generally includes a fan and a core arranged inflow communication with one another. Additionally, the core of the gasturbine engine generally includes, in serial flow order, a compressorsection, a combustion section, a turbine section, and an exhaustsection. In operation, air is provided from the fan to an inlet of thecompressor section where one or more axial compressors progressivelycompress the air until it reaches the combustion section. Fuel is mixedwith the compressed air and burned within the combustion section toprovide combustion gases. The combustion gases are routed from thecombustion section to the turbine section. The flow of combustion gasesthrough the turbine section drives the turbine section and is thenrouted through the exhaust section, e.g., to atmosphere.

More particularly, the combustion section includes a combustor having acombustion chamber defined by a combustor liner. Downstream of thecombustor, the turbine section includes one or more stages, for example,each stage may contain a plurality of stationary nozzle airfoils as wellas a plurality of blade airfoils attached to a rotor that is driven bythe flow of combustion gases against the blade airfoils. The turbinesection may have other configurations as well. In any event, a flow pathis defined by an inner boundary and an outer boundary, which both extendfrom the combustor through the stages of the turbine section.

Typically, the inner and outer boundaries defining the flow pathcomprise separate components. For example, an outer liner of thecombustor, a separate outer band of a nozzle portion of a turbine stage,and a separate shroud of a blade portion of the turbine stage usuallydefine at least a portion of the outer boundary of the flow path.However, utilizing separate components to form each of the outerboundary and the inner boundary requires a great number of parts, e.g.,one or more seals may be required at each interface between the separatecomponents to minimize leakage of fluid from the flow path, which canincrease the complexity and weight of the gas turbine engine withouteliminating leakage points between the separate components. Therefore,flow path assemblies may be utilized that have a unitary construction,e.g., a unitary outer boundary structure, where two or more componentsof the outer boundary are integrated into a single piece, and/or aunitary inner boundary structure, where two or more components of theinner boundary are integrated into a single piece.

A unitary construction of the flow path assembly may be furthered byvarious ways of assembling the turbine nozzle airfoils, which also maybe referred to as stator vanes, with the outer boundary structure andthe inner boundary structure. For example, the outer boundary structureand/or the inner boundary structure each may be constructed as a unitarystructure or together may be constructed as a single unitary structure,with the nozzle airfoils inserted and secured during subsequentassembly. As another example, the nozzle airfoils may be integrallyformed with one of the outer boundary structure or the inner boundarystructure and attached to the other boundary structure during subsequentassembly. Separating the nozzle airfoils from the outer and/or innerboundary structures of the flow path assembly may simplifymanufacturing, as well as reduce internal stresses compared to flow pathassemblies comprising nozzle airfoils that are integral with both theouter and inner boundary structures.

Accordingly, improved flow path assemblies would be desirable. Forexample, a flow path assembly utilizing a unitary outer wall to defineits outer boundary and having a plurality of nozzle airfoils received inpockets in the outer and inner boundaries would be beneficial. Asanother example, a flow path assembly utilizing a unitary inner wall todefine its outer boundary and having a plurality of nozzle airfoilsreceived in pockets in the outer and inner boundaries would beadvantageous. Additionally, an inner wall that defines a plurality ofslots for receipt of bayonets or projections from a plurality of nozzleairfoils to secure the nozzle airfoils to the inner wall would beuseful.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present disclosure, a flow pathassembly for a gas turbine engine is provided. The flow path assemblycomprises an inner wall; a unitary outer wall; and a plurality of nozzleairfoils, each nozzle airfoil having an inner end radially opposite anouter end. The unitary outer wall includes a combustor portion extendingthrough a combustion section of the gas turbine engine and a turbineportion extending through at least a first turbine stage of a turbinesection of the gas turbine engine. The combustor portion and the turbineportion are integrally formed as a single unitary structure. Further,the inner wall and the unitary outer wall define a combustor of thecombustion section. Also, the unitary outer wall defines a plurality ofouter pockets, each outer pocket configured for receipt of the outer endof one of the plurality of nozzle airfoils. The inner wall includes aforward segment and an aft segment and defines a plurality of innerpockets such that a portion of each inner pocket is defined by theforward segment and a remaining portion of each inner pocket is definedby the aft segment. Each inner pocket is configured for receipt of theinner end of one of the plurality of nozzle airfoils such that a nozzleairfoil extends from each inner pocket to a respective outer pocket.

In another exemplary embodiment of the present disclosure, a flow pathassembly for a gas turbine engine is provided. The flow path assemblycomprises an inner wall; an outer wall; and a plurality of nozzleairfoils, each nozzle airfoil having an inner end radially opposite anouter end. The inner wall and the unitary outer wall define a combustorof the combustion section. Moreover, the inner wall defines a pluralityof inner pockets, each inner pocket configured for receipt of the innerend of one of the plurality of nozzle airfoils. The outer wall includesa forward segment and an aft segment and defines a plurality of outerpockets such that a portion of each outer pocket is defined by theforward segment and a remaining portion of each outer pocket is definedby the aft segment. Each outer pocket is configured for receipt of theouter end of one of the plurality of nozzle airfoils such that a nozzleairfoil extends from each inner pocket to a respective outer pocket.

In a further exemplary embodiment of the present disclosure, a flow pathassembly for a gas turbine engine is provided. The flow path assemblycomprises an inner wall defining a plurality of bayonet slots and aplurality of recesses along an aft surface of the inner wall. The flowpath assembly also comprises a unitary outer wall including a combustorportion extending through a combustion section of the gas turbine engineand a turbine portion extending through at least a first turbine stageof a turbine section of the gas turbine engine. The turbine portionincludes a plurality of nozzle airfoils, and the combustor portion andthe turbine portion are integrally formed as a single unitary structuresuch that the plurality of nozzle airfoils is integral with the outerwall. The flow path assembly further comprises a first support memberpositioned radially inward of the inner wall to support the inner walland a second support member positioned axially aft of the first supportmember. The second support member includes a plurality of tabs. An innerend of each nozzle airfoil is positioned against the inner wall, andeach tab is received in a recess of the plurality of recesses defined inthe inner wall such that the second support member fits against the aftsurface of the inner wall to cover the plurality of bayonet slots.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 provides a schematic cross-section view of an exemplary gasturbine engine according to various embodiments of the present subjectmatter.

FIG. 2 provides a schematic exploded cross-section view of a combustionsection and a high pressure turbine section of the gas turbine engine ofFIG. 1 according to an exemplary embodiment of the present subjectmatter.

FIG. 3A provides a schematic cross-section view of the combustionsection and high pressure turbine section of FIG. 2 according to anexemplary embodiment of the present subject matter.

FIGS. 3B, 3C, 3D, and 3E provide schematic cross-section views of thecombustion section and high pressure turbine section of FIG. 2 accordingto other exemplary embodiments of the present subject matter.

FIG. 3F provides a partial perspective view of a portion of an integralouter boundary structure and inner boundary structure of the combustionsection and high pressure turbine section of FIG. 2 according to anexemplary embodiment of the present subject matter.

FIG. 4 provides a schematic cross-section view of a portion of a flowpath assembly according to an exemplary embodiment of the presentsubject matter.

FIG. 5 provides a schematic cross-section view of a portion of a flowpath assembly according to another exemplary embodiment of the presentsubject matter.

FIG. 6 provides a schematic cross-section view of the flow path assemblyof FIG. 5 , where the view is radially downward.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention. As used herein, theterms “first,” “second,” and “third” may be used interchangeably todistinguish one component from another and are not intended to signifylocation or importance of the individual components. The terms“upstream” and “downstream” refer to the relative direction with respectto fluid flow in a fluid pathway. For example, “upstream” refers to thedirection from which the fluid flows and “downstream” refers to thedirection to which the fluid flows.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1 , the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” Asshown in FIG. 1 , the turbofan engine 10 defines an axial direction A(extending parallel to a longitudinal centerline 12 provided forreference) and a radial direction R. In general, the turbofan 10includes a fan section 14 and a core turbine engine 16 disposeddownstream from the fan section 14.

The exemplary core turbine engine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure(HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HPcompressor 24. A low pressure (LP) shaft or spool 36 drivingly connectsthe LP turbine 30 to the LP compressor 22. In other embodiments ofturbofan engine 10, additional spools may be provided such that engine10 may be described as a multi-spool engine.

For the depicted embodiment, fan section 14 includes a fan 38 having aplurality of fan blades 40 coupled to a disk 42 in a spaced apartmanner. As depicted, fan blades 40 extend outward from disk 42 generallyalong the radial direction R. The fan blades 40 and disk 42 are togetherrotatable about the longitudinal axis 12 by LP shaft 36. In someembodiments, a power gear box having a plurality of gears may beincluded for stepping down the rotational speed of the LP shaft 36 to amore efficient rotational fan speed.

Referring still to the exemplary embodiment of FIG. 1 , disk 42 iscovered by rotatable front nacelle 48 aerodynamically contoured topromote an airflow through the plurality of fan blades 40. Additionally,the exemplary fan section 14 includes an annular fan casing or outernacelle 50 that circumferentially surrounds the fan 38 and/or at least aportion of the core turbine engine 16. It should be appreciated thatnacelle 50 may be configured to be supported relative to the coreturbine engine 16 by a plurality of circumferentially-spaced outletguide vanes 52. Moreover, a downstream section 54 of the nacelle 50 mayextend over an outer portion of the core turbine engine 16 so as todefine a bypass airflow passage 56 therebetween.

During operation of the turbofan engine 10, a volume of air 58 entersturbofan 10 through an associated inlet 60 of the nacelle 50 and/or fansection 14. As the volume of air 58 passes across fan blades 40, a firstportion of the air 58 as indicated by arrows 62 is directed or routedinto the bypass airflow passage 56 and a second portion of the air 58 asindicated by arrows 64 is directed or routed into the LP compressor 22.The ratio between the first portion of air 62 and the second portion ofair 64 is commonly known as a bypass ratio. The pressure of the secondportion of air 64 is then increased as it is routed through the highpressure (HP) compressor 24 and into the combustion section 26, where itis mixed with fuel and burned to provide combustion gases 66.

The combustion gases 66 are routed through the HP turbine 28 where aportion of thermal and/or kinetic energy from the combustion gases 66 isextracted via sequential stages of HP turbine stator vanes 68 that arecoupled to the outer casing 18 and HP turbine rotor blades 70 that arecoupled to the HP shaft or spool 34, thus causing the HP shaft or spool34 to rotate, thereby supporting operation of the HP compressor 24. Thecombustion gases 66 are then routed through the LP turbine 30 where asecond portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LP turbine stator vanes 72that are coupled to the outer casing 18 and LP turbine rotor blades 74that are coupled to the LP shaft or spool 36, thus causing the LP shaftor spool 36 to rotate, thereby supporting operation of the LP compressor22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the core turbine engine 16 to provide propulsivethrust. Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section 76 of the turbofan 10, also providing propulsive thrust.The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through the core turbine engine 16.

It will be appreciated that, although described with respect to turbofan10 having core turbine engine 16, the present subject matter may beapplicable to other types of turbomachinery. For example, the presentsubject matter may be suitable for use with or in turboprops,turboshafts, turbojets, industrial and marine gas turbine engines,and/or auxiliary power units.

In some embodiments, components of turbofan engine 10, particularlycomponents within hot gas path 78, such as components of combustionsection 26, HP turbine 28, and/or LP turbine 30, may comprise a ceramicmatrix composite (CMC) material, which is a non-metallic material havinghigh temperature capability. Of course, other components of turbofanengine 10, such as components of HP compressor 24, may comprise a CMCmaterial. Exemplary CMC materials utilized for such components mayinclude silicon carbide (SiC), silicon, silica, or alumina matrixmaterials and combinations thereof. Ceramic fibers may be embeddedwithin the matrix, such as oxidation stable reinforcing fibers includingmonofilaments like sapphire and silicon carbide (e.g., Textron's SCS-6),as well as rovings and yarn including silicon carbide (e.g., NipponCarbon's NICALON®, Ube Industries' TYRANNO®, and Dow Corning'sSYLRAMIC®), alumina silicates (e.g., Nextel's 440 and 480), and choppedwhiskers and fibers (e.g., Nextel's 440 and SAFFIL®), and optionallyceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinationsthereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica,talc, kyanite, and montmorillonite). For example, in certainembodiments, bundles of the fibers, which may include a ceramicrefractory material coating, are formed as a reinforced tape, such as aunidirectional reinforced tape. A plurality of the tapes may be laid uptogether (e.g., as plies) to form a preform component. The bundles offibers may be impregnated with a slurry composition prior to forming thepreform or after formation of the preform. The preform may then undergothermal processing, such as a cure or burn-out to yield a high charresidue in the preform, and subsequent chemical processing, such asmelt-infiltration or chemical vapor infiltration with silicon, to arriveat a component formed of a CMC material having a desired chemicalcomposition. In other embodiments, the CMC material may be formed as,e.g., a carbon fiber cloth rather than as a tape.

As stated, components comprising a CMC material may be used within thehot gas path 78, such as within the combustion and/or turbine sectionsof engine 10. As an example, the combustion section 26 may include acombustor formed from a CMC material and/or one or more stages of one ormore stages of the HP turbine 28 may be formed from a CMC material.However, CMC components may be used in other sections as well, such asthe compressor and/or fan sections. Of course, in some embodiments,other high temperature materials and/or other composite materials may beused to form one or more components of engine 10.

FIG. 2 provides an exploded view of a schematic cross-section of thecombustion section 26 and the HP turbine 28 of the turbine section ofthe turbofan engine 10 according to an exemplary embodiment of thepresent subject matter. FIG. 3A provides an unexploded schematiccross-sectional view of the combustion section 26 and the HP turbine 28of FIG. 2 that focuses on an outer boundary of a flow path through thecombustion section 26 and HP turbine 28. The depicted combustion section26 includes a generally annular combustor 80, and downstream of thecombustion section 26, the HP turbine 28 includes a plurality of turbinestages. More particularly, for the depicted embodiment, HP turbine 28includes a first turbine stage 82 and a second turbine stage 84. Inother embodiments, the HP turbine 28 may comprise a different number ofturbine stages; for example, the HP turbine 28 may include one turbinestage or more than two turbine stages. The first turbine stage 82 ispositioned immediately downstream of the combustion section 26, and thesecond turbine stage 84 is positioned immediately downstream of thefirst turbine stage 82. Further, each turbine stage 82, 84 comprises anozzle portion and a blade portion; the first turbine stage 82 includesnozzle portion 82N and blade portion 82B, and the second turbine stage84 includes nozzle portion 84N and blade portion 84B. The nozzle portion82N of the first turbine stage 82 is located immediately downstream ofthe combustion section 26, such that the nozzle portion 82N of the firstturbine stage 82 also may be referred to as a combustor dischargenozzle. Moreover, combustor 80 defines a generally annular combustionchamber 86 such that the combustor 80 may be described as a generallyannular combustor.

Additionally, as described in greater detail below, a flow path 100through the combustion section 26 and the HP turbine 28 is defined by anouter boundary and an inner boundary of a flow path assembly 101. Theouter and inner boundaries form a flow path for the combustion gases 66through the combustion section 26 and HP turbine 28; thus, the flow path100 may comprise at least a portion of the hot gas path 78 describedabove. Further, in other embodiments, the flow path 100 also may extendthrough LP turbine 30 and jet exhaust 32; in still other embodiments,the flow path 100 also may extend forward upstream of the combustionsection 26, e.g., into HP compressor 24. As such, it will be appreciatedthat the discussion herein of the present subject matter with respect tocombustion section 26 and HP turbine 28 is by way of example only andalso may apply to different configurations of gas turbine engines andflow paths 100.

As shown in the exploded view of FIG. 2 , the outer and inner boundariesmay be defined by an outer wall 102 and an inner wall 120, respectively,which may include several portions of the combustion section 26 and HPturbine 28. For instance, the combustor 80 includes an outer liner 108defining an outer boundary of the flow path through the combustor 80.Each nozzle portion 82N, 84N comprises an outer band defining an outerboundary of a flow path through the nozzle portion of each turbinestage, and each blade portion 82B, 84B comprises a shroud defining anouter boundary of a flow path through the blade portion of each turbinestage. More particularly, as shown in FIG. 2 , the first turbine stagenozzle portion 82N comprises outer band 110, first turbine stage bladeportion 82B comprises shroud 112, second turbine stage nozzle portion84N comprises outer band 114, and second turbine stage blade portion 84Bcomprises shroud 116. These portions of the combustion section 26 and HPturbine 28 may comprise at least a portion of the outer wall 102, asdescribed in greater detail below.

Further, as illustrated in FIG. 2 , the combustor 80 includes an innerliner 122 defining an inner boundary of the flow path through thecombustor 80. Each nozzle portion 82N, 84N comprises an inner banddefining an inner boundary of the flow path through the nozzle portionof each turbine stage, and each blade portion 82B, 84B comprises one ormore blade platforms that define an inner boundary of the flow paththrough the blade portion of each turbine stage. More particularly, asshown in FIG. 2 , the first turbine stage nozzle portion 82N comprisesinner band 124, first turbine stage blade portion 82B comprises bladeplatforms 132, second turbine stage nozzle portion 84N comprises innerband 136, and second turbine stage blade portion 84B comprises bladeplatforms 132. These portions of the combustion section 26 and HPturbine 28 may comprise at least a portion of the inner wall 122, asdescribed in greater detail below.

Moreover, in the depicted embodiment, a combustor dome 118 extendsradially across a forward end 88 of the combustor 80. The combustor dome118 may be a part of outer wall 102, may be a part of inner wall 120,may be a part of both outer wall 102 and inner wall 120 (e.g., a portionof the combustor dome 118 may be defined by the outer wall 102 and theremainder may be defined by the inner wall 120), or may be a separatecomponent from outer wall 102 and inner wall 120. Additionally, aplurality of nozzle airfoils is positioned in each of the nozzleportions 82N, 84N. Each nozzle airfoil 126 within the first turbinestage nozzle portion 82N extends radially from the outer band 110 to theinner band 124, and the nozzle airfoils 126 are spaced circumferentiallyabout the longitudinal centerline 12. Each nozzle airfoil 128 within thesecond turbine stage nozzle portion 84N extends radially from the outerband 114 to the inner band 136, and the nozzle airfoils 128 are spacedcircumferentially about the longitudinal centerline 12. Further, aplurality of blade airfoils 130 are positioned in each of the bladeportions 82B, 84B. Each blade airfoil 130 within the first turbine stageblade portion 82B is attached to blade platform 132, which in turn isattached to a first stage rotor 134. The blade airfoils 130 attached tothe first stage rotor 134 are spaced circumferentially about thelongitudinal centerline 12. Similarly, each blade airfoil 130 within thesecond turbine stage blade portion 84B is attached to a blade platform132, which in turn is attached to a second stage rotor 138. The bladeairfoils 130 attached to the second stage rotor 138 are spacedcircumferentially about the longitudinal centerline 12. Each bladeairfoils 130 extends radially outward toward the outer wall 102, i.e.,the outer boundary of the flow path 100, and a clearance gap is definedbetween a tip 140 of each blade airfoil 130 and the outer wall 102 suchthat each turbine rotor 134, 138 is free to rotate within its respectiveturbine stage. Although not depicted, each turbine rotor 134, 138 of theHP turbine 28 is connected to the HP shaft 34 (FIG. 1 ). In such manner,rotor blade airfoils 130 may extract kinetic energy from the flow ofcombustion gases through the flow path 100 defined by the HP turbine 28as rotational energy applied to the HP shaft 34.

Accordingly, flow path 100 through the combustion section 26 and the HPturbine 28 is defined by a flow path assembly 101 having an innerboundary and an outer boundary, and the inner and outer boundariesdefine the flow path for the combustion gases 66 through the combustionsection 26 and HP turbine 28. Portions of the outer boundary of the flowpath assembly 101 may be integrated or unified into a single piece outerwall 102 that defines the radially outer boundary of the gas flow path100. For instance, the outer wall 102 may include a combustor portion104 extending through a combustion section, such as combustion section26, and a turbine portion 106 extending through at least a first turbinestage of a turbine section, such as first turbine stage 82 of HP turbine28. The combustor portion 104 and turbine portion 106 are integrallyformed such that the combustor portion and the turbine portion are asingle unitary structure, i.e., a unitary outer wall 102.

In the exemplary embodiment depicted in FIG. 3A, the outer wall 102includes a combustor portion 104 extending through the combustionsection 26 and a turbine portion 106 extending through at least thefirst turbine stage 82 and the second turbine stage 84 of the turbinesection. In other embodiments, the turbine portion 106 may extendthrough fewer stages (e.g., through one turbine stage as just described)or through more stages (e.g., through one or more stages of the LPturbine 30 positioned downstream of HP turbine 28). The combustorportion 104 and the turbine portion 106 are integrally formed such thatthe combustor portion 104 and the turbine portion 106 are a singleunitary structure, which is referred to herein as unitary outer wall102.

The term “unitary” as used herein denotes that the associated component,such as the outer wall 102, is made as a single piece duringmanufacturing, i.e., the final unitary component is a single piece.Thus, a unitary component has a construction in which the integratedportions are inseparable and is different from a component comprising aplurality of separate component pieces that have been joined togetherand, once joined, are referred to as a single component even though thecomponent pieces remain distinct and the single component is notinseparable (i.e., the pieces may be re-separated). The final unitarycomponent may comprise a substantially continuous piece of material, orin other embodiments, may comprise a plurality of portions that arepermanently bonded to one another. In any event, the various portionsforming a unitary component are integrated with one another such thatthe unitary component is a single piece with inseparable portions.

As shown in FIG. 3A, the combustor portion 104 of the unitary structureforming outer wall 102 includes the outer liner 108 of the combustor 80.The turbine portion 106 includes the outer band 110 of the first turbinestage nozzle portion 82N, the shroud 112 of the first turbine stageblade portion 82B, the outer band 114 of the second turbine stage nozzleportion 84N, and the shroud 116 of the second turbine stage bladeportion 84B. As stated, these outer boundary components are integratedinto a single piece to form the unitary structure that is outer wall102. Thus, in the exemplary embodiment of FIG. 2 , outer liner 108,outer band 110, shroud 112, outer band 114, and shroud 116 areintegrally formed, i.e., constructed as a single unit or piece to formthe integrated or unitary outer wall 102.

In some embodiments, other portions of the flow path assembly 101 may beintegrated into the unitary structure of outer wall 102, and in stillother embodiments, at least a portion of the outer boundary and theinner boundary are made as a single, unitary component such that theflow path assembly 101 may be referred to as an integrated flow pathassembly. For example, referring to FIG. 3B, the combustor portion 104of unitary outer wall 102 also may include the combustor dome 118 thatextends across the forward end 88 of combustor 80. As such, in theexemplary embodiment of FIG. 3B, the outer liner 108, outer band 110,shroud 112, outer band 114, shroud 116, and combustor dome 118 areconstructed as a single unit or piece to form the integrated or unitaryouter wall 102. That is, the outer liner 108, outer bands 110, 114,shrouds 112, 116, and combustor dome 118 are integrally formed such thatthe outer liner 108, outer bands 110, 114, shrouds 112, 116, andcombustor dome 118 are a single unitary structure.

As another example, referring to FIG. 3C, at least a portion of theinner wall 120 defining the inner boundary of the flow path 100 may beintegrated with the outer wall 102 to form an integrated flow pathassembly 101. In the exemplary embodiment of FIG. 3C, the combustorportion 104 further comprises the inner liner 122, such that the innerliner 122 is integrated with the unitary structure of the outer wall 102shown in FIG. 3B. Thus, the outer liner 108, outer band 110, shroud 112,outer band 114, shroud 116, combustor dome 118, and inner liner 122 areintegrally formed such that the outer liner 108, outer bands 110, 114,shrouds 112, 116, combustor dome 118, and inner liner 122 are a singleunitary structure. In the exemplary embodiment of FIG. 3D, the turbineportion 106 further includes the inner band 124 of the first turbinestage nozzle portion 82N, such that the inner band 124 is integratedwith the unitary structure of the flow path assembly 101 shown in FIG.3C. Accordingly, the outer liner 108, outer band 110, shroud 112, outerband 114, shroud 116, combustor dome 118, inner liner 122, and innerband 124 are integrally formed such that the outer liner 108, outerbands 110, 114, shrouds 112, 116, combustor dome 118, inner liner 122,and inner band 124 are a single unitary structure. In the exemplaryembodiment of FIG. 3E, the turbine portion 106 further includes theplurality of nozzle airfoils 126, such that each nozzle airfoil 126 ofthe plurality of nozzle airfoils 126 of the first turbine stage nozzleportion 82N is integrated with the unitary structure of the flow pathassembly 101 shown in FIG. 3D. Therefore, the outer liner 108, outerband 110, shroud 112, outer band 114, shroud 116, combustor dome 118,inner liner 122, inner band 124, and nozzle airfoils 126 are integrallyformed such that the outer liner 108, outer bands 110, 114, shrouds 112,116, combustor dome 118, inner liner 122, inner band 124, and nozzleairfoils 126 are a single unitary structure.

Of course, the nozzle airfoils 126 of the first turbine stage nozzleportion 82N may be integrated with the outer wall 102 without beingintegrated with the inner wall 120. For example, the plurality of nozzleairfoils 126 may be formed as a single unit or piece with the outerliner 108, outer band 110, shroud 112, outer band 114, shroud 116 suchthat the outer liner 108, outer bands 110, 114, shrouds 112, 116, andnozzle airfoils 126 are a single unitary structure, i.e., a unitaryouter wall 102. In other embodiments, the unitary outer wall 102 alsomay include the combustor dome 118, such that the outer liner 108, outerband 110, shroud 112, outer band 114, shroud 116, combustor dome 118,and nozzle airfoils 126 are integrally formed or constructed as a singleunit or piece. In still other embodiments, the inner liner 122 also maybe included, such that the outer liner 108, outer band 110, shroud 112,outer band 114, shroud 116, combustor dome 118, inner liner 122, andnozzle airfoils 126 are integrally formed as a single unitary structure,i.e., a unitary outer wall 102.

In yet other embodiments, the combustor dome 118 may not be integratedwith either the outer wall 102 or the inner wall 120 in whole or inpart. That is, the combustor dome 118 is a separate component from boththe outer wall 102 and the inner wall 120. As such, the flow path 100may be discontinuous between the combustor dome 118 and outer wall 102,as well as between combustor dome 118 and inner wall 120. Further, insuch embodiments, the combustor dome 118 is configured to move axiallywith respect to the inner wall 120 and the outer wall 102 but may beattached to, and accordingly supported by, one or more fuel nozzleassemblies 90. More particularly, an axial slip joint may be formedbetween the combustor dome 118 and each of the outer wall 102 and theinner wall 120 such that the combustor dome 118 may move or floataxially with respect to the inner wall 120 and outer wall 102. Allowingthe combustor dome 118 to float relative to the outer wall 102 and innerwall 120 can help control the position of the fuel nozzle assembly 90with respect to the combustor dome 118 and combustor 80. For example,the combustor dome 118, outer wall 102, and inner wall 120 may be madeof a different material or materials than the fuel nozzle assembly 90.As described in greater detail below, in an exemplary embodiment, thecombustor dome 118, outer wall 102, and inner wall 120 are made from aceramic matrix composite (CMC) material, and the fuel nozzle assembly 90may be made from a metallic material, e.g., a metal alloy or the like.In such embodiment, the CMC material thermally grows or expands at adifferent rate than the metallic material. Thus, allowing the combustordome 118 to move axially with respect to outer and inner walls 102, 120may allow for tighter control of the immersion of swirler 92 of fuelnozzle assembly 90 within combustor dome 118, as well as combustor 80,than if the combustor dome 118 was attached to the outer and inner walls102, 120. Tighter control of the position of fuel nozzle assembly 90 andits components with respect to combustor 80 can reduce variation inoperability and performance of engine 10.

Additionally, in embodiments in which the combustor dome 118 is separatefrom the outer and inner walls 102, 120, the outer wall 102 and innerwall 120 also may move axially and radially with respect to thecombustor dome 118. By decoupling the combustor dome 118 from the walls102, 120 and allowing relative movement between the walls 102, 120 andthe combustor dome 118, stress coupling may be alleviated between theouter and inner walls 102, 120 and the combustor dome 118. Moreover, anyleakage between the uncoupled combustor dome 118 and outer and innerwalls 102, 120 may be utilized as purge and/or film starter flow.

FIG. 3F provides a partial perspective view of a portion of an integralflow path assembly 101, having an outer wall 102 and inner wall 120formed as a single piece component. As described with respect to FIG. 3Dand shown in FIG. 3F, in some embodiments of the combustion gas flowpath assembly 101, the outer liner 108, outer band 110, shroud 112,outer band 114, shroud 116, combustor dome 118, inner liner 122, andinner band 124 are integrally formed such that the outer liner 108,outer bands 110, 114, shrouds 112, 116, combustor dome 118, inner liner122, and inner band 124 are a single unitary structure. FIG. 3F furtherillustrates that a plurality of openings 142 for receipt of fuel nozzleassemblies 90 and/or swirlers 92 may be defined in the forward end 88 ofcombustor 80 of the unitary flow path assembly 101. Further, it will beappreciated that FIG. 3F illustrates only a portion of the integral flowpath assembly 101 and that, although its entire circumference is notillustrated in FIG. 3F, the flow path assembly 101 is a single, unitarypiece circumferentially as well as axially. As such, the integral flowpath assembly 101 defines a generally annular, i.e., generallyring-shaped, flow path between the outer wall 102 and inner wall 120.

Integrating various components of the outer and inner boundaries of theflow path assembly 101 as described above can reduce the number ofseparate pieces or components within engine 10, as well as reduce theweight, leakage, and complexity of the engine 10, compared to known gasturbine engines. For instance, known gas turbine engines employ seals orsealing mechanisms at the interfaces between separate pieces of the flowpath assembly to attempt to minimize leakage of combustion gases fromthe flow path. By integrating the outer boundary, for example, asdescribed with respect to unitary outer wall 102, split points orinterfaces between the outer combustor liner and first turbine stageouter band, the first turbine stage outer band and the first turbinestage shroud, etc. can be eliminated, thereby eliminating leakage pointsas well as seals or sealing mechanisms required to prevent leakage.Similarly, by integrating components of the inner boundary, split pointsor interfaces between the integrated inner boundary components areeliminated, thereby eliminating leakage points and seals or sealingmechanisms required at the inner boundary. Accordingly, undesiredleakage, as well as unnecessary weight and complexity, can be avoided byutilizing unitary components in the flow path assembly. Other advantagesof unitary outer wall 102, unitary inner wall 120, and/or a unitary flowpath assembly 101 will be appreciated by those of ordinary skill in theart.

As illustrated in FIGS. 3A through 3F, the outer wall 102 and the innerwall 120 define a generally annular flow path therebetween. That is, theunitary outer wall 102 circumferentially surrounds the inner wall 120;stated differently, the unitary outer wall 102 is a single pieceextending 360° degrees about the inner wall 120, thereby defining agenerally annular or ring-shaped flow path therebetween. As such, thecombustor dome 118, which extends across the forward end 88 of thecombustor 80, is a generally annular combustor dome 118. Further, thecombustor dome 118 defines an opening 142 for receipt of a fuel nozzleassembly 90 positioned at forward end 88. The fuel nozzle assembly 90,e.g., provides combustion chamber 86 with a mixture of fuel andcompressed air from the compressor section, which is combusted withinthe combustion chamber 86 to generate a flow of combustion gases throughthe flow path 100. The fuel nozzle assembly 90 may attach to thecombustor dome 118 or may “float” relative to the combustor dome 118 andthe flow path 100, i.e., the fuel nozzle assembly 90 may not be attachedto the combustor dome 118. In the illustrated embodiments, the fuelnozzle assembly 90 includes a swirler 92, and in some embodiments, theswirler 92 may attach to the combustor dome 118, but alternatively, theswirler 92 may float relative to the combustor dome 118 and flow path100. It will be appreciated that the fuel nozzle assembly 90 or swirler92 may float relative to the combustor dome 118 and flow path 100 alongboth a radial direction R and an axial direction A or only along one orthe other of the radial and axial directions R, A. Further, it will beunderstood that the combustor dome 118 may define a plurality ofopenings 142, each opening receiving a swirler 92 or other portion offuel nozzle assembly 90.

As further illustrated in FIGS. 3A through 3F, the flow path assembly101 generally defines a converging-diverging flow path 100. Moreparticularly, the outer wall 102 and the inner wall 120 define agenerally annular combustion chamber 86, which forms a forward portionof the flow path 100. Moving aft or downstream of combustion chamber 86,the outer wall 102 and inner wall 120 converge toward one another,generally in the region of first turbine stage 82. Continuing downstreamof the first turbine stage 82, the outer wall 102 and inner wall 120then diverge, generally in the region of second turbine stage 84. Theouter wall 102 and inner wall 120 may continue to diverge downstream ofthe second turbine stage 84. In exemplary embodiments, e.g., as shown inFIG. 3A and referring only to the unitary outer wall 102, the firstturbine stage nozzle outer band portion 110 and blade shroud portion 112of the outer wall 102 converge toward the axial centerline 12. Thesecond turbine stage nozzle outer band portion 114 and blade shroudportion 116 of the outer wall 102 diverge away from the axial centerline12. As such, the outer boundary of flow path 100 formed by the unitaryouter wall 102 defines a converging-diverging flow path 100.

Turning to FIG. 4 , a cross-sectional view is provided of a portion ofthe flow path assembly 101 according to an exemplary embodiment of thepresent subject matter. As shown in the depicted embodiment, the flowpath assembly 101 includes an inner wall 120 and a unitary outer wall102. As described above, the unitary outer wall 102 includes a combustorportion 104 that extends through the combustion section 26 and a turbineportion 106 that extends through at least a first turbine stage 82 ofthe turbine section 28; in the embodiment of FIG. 4 , the turbineportion extends through the second turbine stage 84. Further, thecombustor portion 104 and the turbine portion 106 of the outer wall 102are integrally formed as a single unitary structure and, thus, may bereferred to as unitary outer wall 102, and the inner wall 120 and theunitary outer wall 102 define the combustor 80.

More particularly, in the illustrated embodiment, the combustor portion104 of the unitary outer wall 102 comprises the outer liner 108 of thecombustor 80, and the turbine portion 106 comprises the outer band 110of the first turbine stage nozzle portion 82N, the shroud 112 of thefirst turbine stage blade portion 82B, the outer band 114 of the secondturbine stage nozzle portion 84N, and the shroud 116 of the secondturbine stage blade portion 84B. The inner wall 120 also may include aunitary structure that may be referred to as unitary inner wall portion.For example, in FIG. 4 , the inner wall 120 comprises a forward segment120 a and an aft segment 120 b, and the forward segment 120 a is aunitary structure that may be referred to as unitary inner wall portion120 a. The unitary inner wall portion 120 a includes the inner liner 122and a forward portion of the first turbine stage inner band 124, whichare integrally formed as unitary inner wall portion 120 a. The aftsegment 120 b comprises the remaining portion of the inner wall 120,i.e., the remainder of the first turbine stage inner band 124. In someembodiments, as previously described, the unitary outer wall 102 orunitary inner wall portion 120 a also may include the combustor dome118, or the unitary outer wall 102 and the unitary inner wall portion120 a each may include a portion of the combustor dome 118. In stillother embodiments, the outer wall 102, combustor dome 118, and theforward segment 120 a of inner wall 120 may be integrally formed as asingle piece, unitary structure.

Referring still to the embodiment illustrated in FIG. 4 , the flow pathassembly 101 further includes a plurality of first turbine stage nozzleairfoils 126. Each of the plurality of nozzle airfoils 126 extendsbetween pockets defined in the inner boundary structure and the outerboundary structure of flow path assembly 101. More particularly, eachnozzle airfoil 126 has an inner end 126 a radially opposite an outer end126 b. The unitary outer wall 102 defines a plurality of outer pockets170, and each outer pocket 170 is configured for receipt of the outerend 126 b of one of the plurality of first turbine stage nozzle airfoils126. Similarly, the inner wall 120 defines a plurality of inner pockets174. More specifically, as shown in FIG. 4 , a portion of each innerpocket 174 is defined by the forward segment 120 a of inner wall 120 anda remaining portion of each inner pocket 174 is defined by the aftsegment 120 b. Further, each inner pocket 174 is configured for receiptof the inner end 126 a of one of the plurality of first turbine stagenozzle airfoils 126. Accordingly, a nozzle airfoil 126 extends from eachinner pocket 174 to a respective outer pocket 170.

Keeping with FIG. 4 , the plurality of outer pockets 170 are definedalong an area 178 of an inner surface 103 of the unitary outer wall 102,and the unitary outer wall 102 is built up at the area 178 where theplurality of outer pockets 170 are defined. The built up area 178 mayprovide extra thickness in outer wall 102 such that the outer pockets170 may be defined therein, without requiring the extra thicknessthroughout the entire unitary outer wall 102.

Moreover, the forward segment 120 a of the inner wall 120 comprises aforward flange 182, and the aft segment 120 b of the inner wall 120comprises an aft flange 184. Each flange 182, 184 extends radiallyinward from the inner wall 120, i.e., toward the axial centerline 12.The forward flange 182 is positioned adjacent the aft flange 184 whenthe forward segment 120 a and the aft segment 120 b are assembled in theflow path assembly 101. Each flange 182, 184 defines at least one pinaperture 186 for receipt of a pin 188. As such, at least one pin 188secures the forward flange 182 to the aft flange 184, e.g., to hold theforward and aft segments 120 a, 120 b of inner wall 120 in position withrespect to one another.

It will be appreciated that the configuration illustrated in FIG. 4 isby way of example only. For instance, although FIG. 4 depicts only thefirst turbine stage nozzle portion 82N of the flow path assembly 101,the subject matter described with respect to the first turbine stagenozzle portion 82N applies equally to the second turbine stage nozzleportion 84N, and also may apply to any other nozzle portions of theturbine sections of the engine 10. That is, while only the first turbinestage nozzle airfoils 126 and the adjacent portions of outer wall 102and inner wall 120 are shown and described with respect to FIG. 4 , thesecond turbine stage nozzle airfoils 128 and the adjacent portions ofouter wall 102 and inner wall 120 may be similarly configured. Moreparticularly, the outer wall 102 may define a plurality of second outerpockets for the receipt of the outer ends of nozzle airfoils 128, andthe second stage inner band 136 may comprise forward and aft segmentsthat together define a plurality of second inner pockets for receipt ofthe inner ends of the nozzle airfoils 128. The outer wall 102 may bebuilt up in the area in which the second outer pockets are defined,e.g., to provide sufficient thickness for the second outer pockets to bedefined in the outer wall 102. Further, the forward and aft segments ofthe second stage inner band 136 may each include a flange, and theflanges of the forward and aft inner band segments may be boltedtogether, e.g., to attach the segments to one another. However, in otherembodiments, the first turbine stage nozzle portion 82N may beconfigured as shown in FIG. 4 while the second turbine stage nozzleportion 84N, as well as nozzle portions of other turbine stages, may beconfigured differently from the depicted embodiment.

In other exemplary embodiments, the outer wall 102 may comprise forwardand aft segments and the inner wall 120 may be a unitary inner wall 120.That is, some embodiments may utilize essentially an oppositeconfiguration from the configuration illustrated in FIG. 4 . Moreparticularly, the unitary inner wall 120 may comprise the inner liner122 of the combustor 80 and the inner band 124 of the first turbinestage nozzle portion 82N. Further, the outer wall 102 may comprise aforward segment and an aft segment. The forward segment may be a unitarystructure that may be referred to as unitary outer wall portion. Theunitary outer wall portion includes the outer liner 108 and a forwardportion of the first turbine stage outer band 110, which are integrallyformed as unitary outer wall portion. The aft segment comprises theremaining portion of the outer band 110 and may further include thefirst turbine stage shroud 112 and a forward portion of the secondturbine stage outer band 114. A second aft segment comprises at leastthe remainder of the second turbine stage outer band 114 and also maycomprise the second turbine stage shroud 116. In some embodiments, theforward segment of outer wall 102 or unitary inner wall 120 also mayinclude the combustor dome 118, or the outer wall forward segment andunitary inner wall 120 each may include a portion of the combustor dome118. In still other embodiments, the outer wall forward segment,combustor dome 118, and inner wall 120 may be integrally formed as asingle piece, unitary structure.

Moreover, similar to the embodiment shown in FIG. 4 , each of theplurality of nozzle airfoils 126, 128 may extend between pockets definedin the inner boundary structure and the outer boundary structure of flowpath assembly 101. More specifically, the unitary inner wall 120 definesa plurality of inner pockets 174 and a plurality of second innerpockets. Each inner pocket 174 is configured for receipt of the innerend 126 a of one of the plurality of first turbine stage nozzle airfoils126, and each second inner pocket is configured for receipt of the innerend of one of the plurality of second turbine stage nozzle airfoils 128.

Similarly, the outer wall 102 defines a plurality of outer pockets 170.A portion of each outer pocket 170 is defined by the forward segment ofouter wall 102 and a remaining portion of each outer pocket 170 isdefined by the aft segment. A plurality of second outer pockets also maybe defined in the outer wall 102, e.g., a portion of each second outerpocket may be defined by the forward portion of the second turbine stageouter band portion of the outer wall aft segment, and a remainingportion of each second outer pocket may be defined by the second aftsegment, which includes the remainder of the second turbine stage outerband. Further, each outer pocket 170 is configured for receipt of theouter end 126 b of one of the plurality of first turbine stage nozzleairfoils 126, and each second outer pocket is configured for receipt ofthe outer end of one of the plurality of second turbine stage nozzleairfoils 128. Accordingly, a first turbine stage nozzle airfoil 126extends from each inner pocket 174 to a respective outer pocket 170, anda second turbine stage nozzle airfoil 128 extends from each second innerpocket to a respective second outer pocket.

Further, similar to the embodiment shown in FIG. 4 , the plurality ofinner pockets 174 may be defined along an area of an inner surface 121of the unitary inner wall 120, and the unitary inner wall 120 may bebuilt up at the area where the plurality of inner pockets 174 aredefined. Likewise, the plurality of second inner pockets may be definedalong an area of inner surface 121, and the unitary inner wall 120 maybe built up at the area where the plurality of second inner pockets aredefined. The built up areas near the inner pockets may provide extrathickness in inner wall 120 such that the inner pockets may be definedtherein, without requiring the extra thickness throughout the entiretyof inner wall 120.

Moreover, the forward and aft segments of the outer wall 102 maycomprise flanges for securing the outer wall segments to one another.For example, the forward segment of the outer wall 102 may comprise aforward flange, the aft segment of the outer wall 102 may comprise afirst aft flange and a second aft flange, and the second aft segment maycomprise a third aft flange. Each flange may extend radially outwardfrom the outer wall 102, i.e., away from the axial centerline 12. Theforward flange may be positioned adjacent the first aft flange when theforward and aft segments are assembled in the flow path assembly 101,and the second aft flange may be positioned adjacent the third aftflange when the aft segment and the second aft segment are assembled inthe flow path assembly 101. Each flange defines at least one pinaperture 186 for receipt of a pin 188. As such, at least one pin 188secures the forward flange to the first aft flange, e.g., to hold theforward and aft segments of outer wall 102 in position with respect toone another. Similarly, at least one pin 188 secures the second aftflange to the third aft flange, e.g., to hold the aft segment and thesecond aft segment of outer wall 102 in position with respect to oneanother.

It will be understood that each pocket 170, 172, 174, 176 may have ashape substantially similar to an axial cross-sectional shape of therespective airfoil 126, 128 received within the pocket. That is, eachouter pocket 170, 172 may generally be described as an airfoil-shapeddepression in the outer wall 102, and each inner pocket 174, 176 maygenerally be described as an airfoil-shaped depression in the inner wall120 and inner band 136. More specifically, the first outer pockets 170and first inner pockets 174 may have a shape corresponding to the firstturbine stage nozzle airfoils 126, the second outer pockets 172 andsecond inner pockets 176 may have a shape corresponding to the secondturbine stage nozzle airfoils 128. However, the pockets 170, 172, 174,176 may have any appropriate shape for receiving the nozzle airfoils126, 128.

Utilizing an outer wall 102 or an inner wall 120 that comprises forwardand aft segments may help simplify the assembly of the flow pathassembly 101. For example, referring to the exemplary embodiment of FIG.4 , the outer wall 102 and the forward segment 120 a of the inner wall120 may be positioned in the assembly. Next, the plurality of firstturbine stage nozzle airfoils 126 may be installed, e.g., by insertingthe outer ends 126 b of the airfoils 126 into the first outer pockets170 and positioning the inner ends 126 a within the portion of the firstinner pockets 174 defined by the forward segment 120 a of inner wall120. Then, the aft segment 120 b of the inner wall 120 may be slid intoposition, with the remainder of the first inner pockets 174 defined inthe aft segment 120 b enclosing the inner ends 126 a of the firstturbine stage nozzle airfoils 126 in the first inner pockets 174. Thefirst stage rotor 134 may then be installed, followed by the forwardsegment 136 a of the second turbine stage inner band 136. Next, thesecond turbine stage nozzle airfoils 128 may be installed, e.g., byinserting the outer ends 128 b of the airfoils 128 into the second outerpockets 172 and positioning the inner ends 128 a within the portion ofthe second inner pockets 176 defined by the forward segment 136 a ofinner band 136. Then, the aft segment 136 b of the inner band 136 may beslid into position, with the remainder of the second inner pockets 176defined in the aft segment 136 b enclosing the inner ends 128 a of thesecond turbine stage nozzle airfoils 128 in the second inner pockets176. The second stage rotor 138 may be installed to complete theassembly of the flow path assembly 101, as well as the combustionsection 26 and HP turbine section 28 of the engine 10.

Turning now to FIGS. 5 and 6 , another exemplary embodiment of the flowpath assembly 101 is provided. The depicted embodiment utilizes abayonet type joint to secure the nozzle airfoils 126, 128, which areintegral with the outer wall 102, to the inner wall 120 and inner band136. That is, referring particularly to FIG. 5 , the first turbine stagenozzle airfoils 126 are integrated with the outer wall 102 to form aunitary structure with the outer boundary of the flow path assembly 101.Thus, the outer wall 102 depicted in FIG. 5 is similar to the embodimentof FIG. 3E, although the airfoils 126 of FIG. 5 are not integrated withthe inner wall 120, and the inner wall 120 and combustor dome 118 may,but need not, be integral with the unitary outer wall 102.

Accordingly, as shown in FIG. 5 , the inner end 126 a of the nozzleairfoils 126 is positioned against the inner surface 121 of the innerwall 120, and each inner end 126 a includes a bayonet that is receivedin a bayonet slot defined in the inner wall 120 as described in greaterdetail below. Further, the inner wall 120 includes a built up portion orarea 119 immediately upstream or forward of the nozzle airfoils 126,such that the built up area 119 helps shield the interface between thenozzle airfoils 126 and the inner wall 120 from the hot combustiongases. It also will be appreciated that the built up area 119 of innerwall 120 may function similar to fillets that may be defined on nozzleairfoils 126 in other embodiments of the airfoils. The built up area 119may be an area of built up CMC material that forms the inner wall 120.

Keeping with FIG. 5 , the inner end 126 a of each nozzle airfoil 126includes a projection or bayonet 190, which is received in a bayonetslot 192 defined in the inner wall 120. More particularly, referring toFIG. 6 , the inner wall 120 defines a plurality of bayonet slots 192such that each bayonet 190 is received in a bayonet slot 192 of theinner wall 120. Each slot 192 generally has an L-shape, i.e., eachbayonet slot 192 has a first leg 192 a and a second leg 192 b thatextends generally perpendicular to the first leg 192 a. The first andsecond legs 192 a, 192 b are joined together to form a continuouspathway, i.e., bayonet slot 192. Thus, when a bayonet 190 of a nozzleairfoil 126 is received within a bayonet slot 192, the inner wall 120may be slid such that the bayonet 190 moves along the first leg 192 aand then the inner wall 120 may be rotated such that the bayonet 190then moves along the second leg 192 b until the bayonet 190 is seatedwithin an end 192 c of the bayonet slot 192. Thus, the nozzle airfoils126 are secured in the inner wall 120 via bayonet joints, which preventaxial movement of the nozzle airfoil bayonets 190 within the inner wallbayonet slots 192 to secure the airfoils 126 in the inner wall 120.

Further, the bayonet slots 192 are defined such that the slots 192 areoriented to take any tangential and axial loads from the nozzle airfoils126. That is, the bayonet slots 192 are oriented such that a tangentialand/or axial load on a nozzle airfoil 126 loads the airfoil into theinner wall 120 surrounding the bayonet 190 rather than causing thebayonet 190 to move along or within the slot 192. Therefore, the bayonetslots 192 are defined in a manner that prevents the nozzle airfoils 126from being dislodged from or moving from the slot ends 192 c when theairfoils are subject to aerodynamic loads.

As also illustrated in FIGS. 5 and 6 , a first support member 194 and asecond support member 196 are positioned radially inward of the innerwall 120, i.e., closer to the axial centerline 12 of engine 10, tosupport the inner wall. The first support member 194 defines a groove198 for receipt of a projection 200 of the inner wall 120. The innerwall projection 200 provides an area for bayonet slots 192 to be definedin the inner wall 120 for the receipt of airfoil bayonets 190. The firstsupport member groove 198 fits against the projection 200 to support theinner wall 120 at the projection 200. The first support member 194further includes a housing 202 at a forward side 204 of the firstsupport member. The housing 202 is configured for receipt of a seal 206,which preferably is generally ring shaped to extend circumferentiallyagainst the inner wall 120. For example, the seal 206 may be a pistonring seal or other suitable seal. The seal 206 helps prevent air leakageat the forward or upstream end of the inner wall 120.

The second support member 196 is positioned axially aft of the firstsupport member 194. The second support member 196 includes a pluralityof tabs 208. As shown in FIGS. 5 and 6 , the inner wall 120 defines aplurality of recesses 210 along an aft surface 123 of the inner wall.Each tab 208 of the second support member 196 is received in a recess210 of the inner wall 120 to position or fit the second support member196 against the aft surface 123 of inner wall 120 such that the secondsupport member 196 covers the plurality of bayonet slots 192 defined inthe inner wall 120. As such, the second support member 196 helps preventfluid leakage from or into the bayonet slots 192.

As further depicted in FIG. 5 , the first support member defines aflange 212 and the second support member defines a flange 214. Eachflange 212, 214 defines an aperture 216, and a bolt 218 is receivedthrough the aperture 216 of the first support member 194 and theaperture 216 of the second support member 196 to secure the firstsupport member to the second support member. Other ways of securing thefirst and second support members 194, 196 to one another also may beused, and in suitable embodiments, the first and second support members194, 196 may be one integral, single piece component or may be dividedinto more than two components.

It will be appreciated that, although FIGS. 5 and 6 depict only thefirst turbine stage nozzle portion 82N of the flow path assembly 101,the subject matter described with respect to the first turbine stagenozzle portion 82N applies equally to the second turbine stage nozzleportion 84N, and also may apply to any other nozzle portions of theturbine sections of the engine 10. That is, while only the first turbinestage nozzle airfoils 126 are shown and described with respect to FIGS.5 and 6 , the second turbine stage nozzle airfoils 128 may be similarlyconfigured, and the second stage inner band 136 may be configuredsimilarly to the first stage inner band portion of the inner wall 120.Further, first and second support members similar to those depicted inFIGS. 5 and 6 also may be provided in the second turbine stage. However,in other embodiments, the first turbine stage nozzle portion 82N may beconfigured as shown in FIGS. 5 and 6 while the second turbine stagenozzle portion 84N, as well as nozzle portions of other turbine stages,may be configured differently from the depicted embodiment.

As previously stated, the outer wall 102, inner wall 120, and combustordome 118, and in some embodiments, first and second turbine stage nozzleairfoils 126, 128, may comprise a CMC material. More particularly, inexemplary embodiments, the combustor portion 104 and the turbine portion106 of flow path assembly 101 are integrally formed from a CMC materialsuch that the resulting unitary structure is a CMC component. Forexample, where the combustor portion 104 includes the outer liner 108 ofthe combustor 80 and the turbine portion 106 includes the outer band 110of the first turbine stage nozzle portion 82N, the shroud 112 of thefirst turbine stage blade portion 82B, the outer band 114 of the secondturbine stage nozzle portion 84N, and the shroud 116 of the secondturbine stage blade portion 84B, the outer liner 108, outer bands 110,114, and shrouds 114, 116 may be integrally formed from a CMC materialto produce a unitary CMC outer wall 102. As described above, in otherembodiments, additional CMC components, such as the nozzle airfoils 126,128, may be integrally formed with the outer liner 108, outer bands 110,114, and shrouds 114, 116 to construct a unitary CMC outer wall 102.Similarly, the inner wall 120 may be formed from a CMC material. Forinstance, where the inner wall 120 comprises separate components, e.g.,inner liner 122, inner bands 124, 136, and blade platforms 132, eachcomponent of the inner wall 120 may be formed from a CMC material. Inembodiments in which two or more components are integrated to form aunitary inner wall 120, the components may be integrally formed from aCMC material to construct a unitary CMC inner wall 120.

Examples of CMC materials, and particularly SiC/Si—SiC (fiber/matrix)continuous fiber-reinforced ceramic composite (CFCC) materials andprocesses, are described in U.S. Pat. Nos. 5,015,540; 5,330,854;5,336,350; 5,628,938; 6,024,898; 6,258,737; 6,403,158; and 6,503,441,and U.S. Patent Application Publication No. 2004/0067316. Such processesgenerally entail the fabrication of CMCs using multiple pre-impregnated(prepreg) layers, e.g., the ply material may include prepreg materialconsisting of ceramic fibers, woven or braided ceramic fiber cloth, orstacked ceramic fiber tows that has been impregnated with matrixmaterial. In some embodiments, each prepreg layer is in the form of a“tape” comprising the desired ceramic fiber reinforcement material, oneor more precursors of the CMC matrix material, and organic resinbinders. Prepreg tapes can be formed by impregnating the reinforcementmaterial with a slurry that contains the ceramic precursor(s) andbinders. Preferred materials for the precursor will depend on theparticular composition desired for the ceramic matrix of the CMCcomponent, for example, SiC powder and/or one or more carbon-containingmaterials if the desired matrix material is SiC. Notablecarbon-containing materials include carbon black, phenolic resins, andfuranic resins, including furfuryl alcohol (C₄H₃OCH₂OH). Other typicalslurry ingredients include organic binders (for example, polyvinylbutyral (PVB)) that promote the flexibility of prepreg tapes, andsolvents for the binders (for example, toluene and/or methyl isobutylketone (MIBK)) that promote the fluidity of the slurry to enableimpregnation of the fiber reinforcement material. The slurry may furthercontain one or more particulate fillers intended to be present in theceramic matrix of the CMC component, for example, silicon and/or SiCpowders in the case of a Si—SiC matrix. Chopped fibers or whiskers orother materials also may be embedded within the matrix as previouslydescribed. Other compositions and processes for producing compositearticles, and more specifically, other slurry and prepreg tapecompositions, may be used as well, such as, e.g., the processes andcompositions described in U.S. Patent Application Publication No.2013/0157037.

The resulting prepreg tape may be laid-up with other tapes, such that aCMC component formed from the tape comprises multiple laminae, eachlamina derived from an individual prepreg tape. Each lamina contains aceramic fiber reinforcement material encased in a ceramic matrix formed,wholly or in part, by conversion of a ceramic matrix precursor, e.g.,during firing and densification cycles as described more fully below. Insome embodiments, the reinforcement material is in the form ofunidirectional arrays of tows, each tow containing continuous fibers orfilaments. Alternatives to unidirectional arrays of tows may be used aswell. Further, suitable fiber diameters, tow diameters, andcenter-to-center tow spacing will depend on the particular application,the thicknesses of the particular lamina and the tape from which it wasformed, and other factors. As described above, other prepreg materialsor non-prepreg materials may be used as well.

After laying up the tapes or plies to form a layup, the layup isdebulked and, if appropriate, cured while subjected to elevatedpressures and temperatures to produce a preform. The preform is thenheated (fired) in a vacuum or inert atmosphere to decompose the binders,remove the solvents, and convert the precursor to the desired ceramicmatrix material. Due to decomposition of the binders, the result is aporous CMC body that may undergo densification, e.g., melt infiltration(MI), to fill the porosity and yield the CMC component. Specificprocessing techniques and parameters for the above process will dependon the particular composition of the materials. For example, silicon CMCcomponents may be formed from fibrous material that is infiltrated withmolten silicon, e.g., through a process typically referred to as theSilcomp process. Another technique of manufacturing CMC components isthe method known as the slurry cast melt infiltration (MI) process. Inone method of manufacturing using the slurry cast MI method, CMCs areproduced by initially providing plies of balanced two-dimensional (2D)woven cloth comprising silicon carbide (SiC)-containing fibers, havingtwo weave directions at substantially 90° angles to each other, withsubstantially the same number of fibers running in both directions ofthe weave. The term “silicon carbide-containing fiber” refers to a fiberhaving a composition that includes silicon carbide, and preferably issubstantially silicon carbide. For instance, the fiber may have asilicon carbide core surrounded with carbon, or in the reverse, thefiber may have a carbon core surrounded by or encapsulated with siliconcarbide.

Other techniques for forming CMC components include polymer infiltrationand pyrolysis (PIP) and oxide/oxide processes. In PIP processes, siliconcarbide fiber preforms are infiltrated with a preceramic polymer, suchas polysilazane and then heat treated to form a SiC matrix. Inoxide/oxide processing, aluminum or alumino-silicate fibers may bepre-impregnated and then laminated into a preselected geometry.Components may also be fabricated from a carbon fiber reinforced siliconcarbide matrix (C/SiC) CMC. The C/SiC processing includes a carbonfibrous preform laid up on a tool in the preselected geometry. Asutilized in the slurry cast method for SiC/SiC, the tool is made up ofgraphite material. The fibrous preform is supported by the toolingduring a chemical vapor infiltration process at about 1200° C., wherebythe C/SiC CMC component is formed. In still other embodiments, 2D, 2.5D,and/or 3D preforms may be utilized in MI, CVI, PIP, or other processes.For example, cut layers of 2D woven fabrics may be stacked inalternating weave directions as described above, or filaments may bewound or braided and combined with 3D weaving, stitching, or needling toform 2.5D or 3D preforms having multiaxial fiber architectures. Otherways of forming 2.5D or 3D preforms, e.g., using other weaving orbraiding methods or utilizing 2D fabrics, may be used as well.

Thus, a variety of processes may be used to form a unitary structure,such as the outer wall 102 depicted in FIG. 3A, as a unitary CMCcomponent. More specifically, a plurality of plies of a CMC material maybe used to form each unitary structure. The plurality of plies may beinterspersed with one another to integrate the various portions formingthe unitary structure. As an example, the unitary outer wall 102 of FIG.3A may be made from a plurality of outer liner plies, a plurality offirst turbine stage outer band plies, a plurality of first turbine stageshroud plies, a plurality of second turbine stage outer band plies, anda plurality of second turbine stage shroud plies. Where the outer linerplies meet the first turbine stage outer band plies, ends of the outerliner plies may be alternated with ends of the outer band plies tointegrate the plies for forming the outer liner portion with the pliesfor forming the first turbine stage outer band portion of the unitaryouter wall 102. That is, any joints between the plies forming unitaryouter wall 102 may be formed by alternating plies on one side of thejoint with plies on the other side of the joint. As such, the plies forforming unitary outer wall 102 may be interspersed to integrate theplies and, thereby, each portion of the unitary outer wall 102. Ofcourse, the CMC plies may be laid up in other ways as well to form theunitary structure. In addition, laying up the plurality of CMC plies mayinclude defining features of the unitary structure or other component(e.g., inner liner 122 when not integrated with inner band 124 to from aunitary inner wall 120 or separate combustor dome 118) such as openings142 in combustor forward end 88.

After the plurality of CMC plies are laid up to define a unitary CMCcomponent preform, the preform is cured to produce a single piece,unitary CMC component, which is then fired and subjected todensification, e.g., silicon melt-infiltration, to form a final unitaryCMC structure. Continuing with the above outer wall 102 example, theouter wall preform may be processed in an autoclave to produce a greenstate unitary outer wall 102. Then, the green state unitary outer wall102 may be placed in a furnace to burn out excess binders or the likeand then placed in a furnace with a piece or slab of silicon and firedto melt infiltrate the unitary outer wall 102 with at least silicon.More particularly, for unitary outer wall 102 formed from CMC plies ofprepreg tapes that are produced as described above, heating (i.e.,firing) the green state component in a vacuum or inert atmospheredecomposes the binders, removes the solvents, and converts the precursorto the desired ceramic matrix material. The decomposition of the bindersresults in a porous CMC body; the body may undergo densification, e.g.,melt infiltration (MI), to fill the porosity. In the foregoing examplewhere the green state unitary outer wall 102 is fired with silicon, theouter wall 102 undergoes silicon melt-infiltration. However,densification may be performed using any known densification techniqueincluding, but not limited to, Silcomp, melt infiltration (MI), chemicalvapor infiltration (CVI), polymer infiltration and pyrolysis (PIP), andoxide/oxide processes, and with any suitable materials including but notlimited to silicon. In one embodiment, densification and firing may beconducted in a vacuum furnace or an inert atmosphere having anestablished atmosphere at temperatures above 1200° C. to allow siliconor other appropriate material or combination of materials tomelt-infiltrate into the component. The densified CMC body hardens to afinal unitary CMC outer wall 102. In some embodiments, the final unitarystructure may be finish machined, e.g., to bring the structure withintolerance or to define openings 142 in forward end 88, and/or anenvironmental barrier coating (EBC) may be applied to the unitarystructure, e.g., to protect the unitary structure from the hotcombustion gases 66. It will be appreciated that other methods orprocesses of forming CMC components, such as unitary CMC outer wall 102,unitary CMC inner wall 120, or the like may be used as well.

Additionally or alternatively, other processes for producing unitarycomponents may be used to form unitary outer wall 102 and/or unitaryinner wall 120, and the unitary structure(s) may be formed from othermaterials. In some embodiments, an additive manufacturing process may beused to form unitary outer wall 102 and/or unitary inner wall 120. Forexample, an additive process such as Fused Deposition Modeling (FDM),Selective Laser Sintering (SLS), Stereolithography (SLA), Digital LightProcessing (DLP), Direct Metal Laser Sintering (DMLS), Laser Net ShapeManufacturing (LNSM), electron beam sintering or other known process maybe used to produce a unitary outer wall 102 and/or a unitary inner wall120. Generally, an additive process fabricates components usingthree-dimensional information, for example, a three-dimensional computermodel, of the component. The three-dimensional information is convertedinto a plurality of slices, each slice defining a cross section of thecomponent for a predetermined height of the slice. The component is then“built-up” slice by slice, or layer by layer, until finished. Superalloymetallic materials or other suitable materials may be used in anadditive process to form unitary outer wall 102 and/or a unitary innerwall 120. In other embodiments, a unitary outer wall 102 and/or unitaryinner wall 120 may be formed using a forging or casting process. Othersuitable processes or methods may be used as well.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal language of the claims.

What is claimed is:
 1. A flow path assembly for a gas turbine engine,the flow path assembly comprising: an inner wall defining a plurality ofbayonet slots, the inner wall further defining a plurality of recessesalong an aft surface of the inner wall; an outer wall; a plurality ofairfoils extending from the outer wall to the inner wall; a firstsupport member positioned inward of the inner wall along a radialdirection to support the inner wall; and a second support memberpositioned aft of the first support member along an axial direction, thesecond support member including a plurality of tabs, wherein an innerend of each airfoil of the plurality of airfoils is positioned againstthe inner wall, and wherein each tab of the plurality of tabs isreceived in a respective recess of the plurality of recesses such thatthe second support member fits against the aft surface of the innerwall.
 2. The flow path assembly of claim 1, wherein the inner end ofeach airfoil of the plurality of airfoils includes a bayonet that isreceived in a bayonet slot of the plurality of bayonet slots defined inthe inner wall.
 3. The flow path assembly of claim 1, wherein eachbayonet slot of the plurality of bayonet slots includes a first leg thatis perpendicular to a second leg, and wherein the first leg and thesecond leg are joined to define a continuous pathway.
 4. The flow pathassembly of claim 3, wherein the first leg is defined along the axialdirection and the second leg is defined along a circumferentialdirection.
 5. The flow path assembly of claim 1, wherein a respectiverecess of the plurality of recesses is defined between a respective pairof bayonet slots of the plurality of bayonet slots.
 6. The flow pathassembly of claim 1, wherein each bayonet slot of the plurality ofbayonet slots is open along the aft surface of the inner wall, andwherein the second support member fits against the aft surface of theinner wall to cover the plurality of bayonet slots along the aft surfaceof the inner wall.
 7. The flow path assembly of claim 1, wherein thefirst support member includes a housing defined on a forward side of thefirst support member, and wherein a seal is positioned in the housing.8. The flow path assembly of claim 7, wherein the seal is ring shaped,and wherein the seal extends along a circumferential direction againstthe inner wall.
 9. The flow path assembly of claim 1, wherein the firstsupport member defines a flange and the second support member defines aflange, each flange defining an aperture, and wherein a bolt is receivedthrough the aperture of the first support member and the aperture of thesecond support member to secure the first support member to the secondsupport member.
 10. The flow path assembly of claim 1, wherein the innerwall includes a projection extending inward along the radial direction,a depth of each bayonet slot of the plurality of bayonet slots extendingwithin the projection.
 11. The flow path assembly of claim 10, whereinthe first support member defines a groove for receipt of the projection.12. The flow path assembly of claim 1, wherein the inner wall includesan area of built up material immediately upstream of the plurality ofairfoils.
 13. The flow path assembly of claim 12, wherein the inner endof each airfoil of the plurality of airfoils defines an inner endsurface, the inner end surface positioned against an inner surface ofthe inner wall, wherein a bayonet extends from the inner end surface ofeach airfoil of the plurality of airfoils, the bayonet received in arespective bayonet slot of the plurality of bayonet slots defined in theinner wall, and wherein the area of built up material extends outwardalong the radial direction away from the inner end surface of eachairfoil of the plurality of airfoils such that the area of built upmaterial extends radially along a portion of a leading edge of eachairfoil of the plurality of airfoils.
 14. The flow path assembly ofclaim 1, wherein the outer wall includes a combustor portion extendingthrough a combustion section of the gas turbine engine and a turbineportion extending through at least a first turbine stage of a turbinesection of the gas turbine engine, wherein the combustor portion and theturbine portion are integrally formed as a single unitary structure. 15.The flow path assembly of claim 1, wherein the plurality of airfoils areintegrally formed with the outer wall.
 16. A flow path assembly for agas turbine engine, the flow path assembly comprising: an inner walldefining a bayonet slot; an outer wall; a first support memberpositioned inward of the inner wall along a radial direction to supportthe inner wall; and a second support member positioned aft of the firstsupport member along an axial direction, the second support memberincluding a tab; and an airfoil extending from the outer wall to theinner wall, the airfoil having an inner end, a bayonet extending fromthe inner end, wherein the bayonet is received in the bayonet slot,wherein the bayonet slot includes a first leg that is perpendicular to asecond leg, the first leg joining the second leg to define a continuouspathway, and wherein the first leg defines an open end at an aft surfaceof the inner wall, wherein the inner wall defines a recess along the aftsurface of the inner wall, and wherein the tab is received in the recesssuch that the second support member fits against the aft surface of theinner wall and covers the open end of the first leg of the bayonet slot.17. A method for forming a flow path assembly, the flow path assemblydefining an axial direction, a radial direction, and a circumferentialdirection that are perpendicular to one another, the method comprising:sliding an inner wall along the axial direction such that a bayonetextending along the radial direction from an inner end of an airfoil isreceived in and moves along a first leg of a bayonet slot defined in theinner wall, the first leg extending along the axial direction; androtating the inner wall along the circumferential direction such thatthe bayonet moves along a second leg of the bayonet slot, the second legextending along the circumferential direction; and positioning a supportmember against an aft surface of the inner wall such that a tab of thesupport member is received in a recess defined by the inner wall, thesupport member covering an open end of the bayonet slot.
 18. The methodof claim 17, wherein the support member includes a housing defined on aforward side of the support member.
 19. The method of claim 18, whereina seal is positioned in the housing.
 20. The method of claim 17, whereinthe support member defines a flange, and wherein the flange defines anaperture.